Fluid distribution and injection systems

ABSTRACT

Apparatus for fluid distribution and injection within a propulsive device, wherein a plurality of preformed laminations are stacked and bonded together to form an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, the interior surface having discharge pores communicating with corresponding fluid discharge channels within the shell.

United States Patent 1 Pearson [451 Jan. 16, 1973 54 FLUID DISTRIBUTION AND 2,705,399 4 1955 INJECTION SYSTEMS 3,286,474 1/1966 3,309,026 3/1967 [76] Inventor: Ronald K. Pearson, i0350 Vacco 3,353,359 il/l967 St., Hacienda Heights, Calif. 91745 7 I Primary ExaminerDouglas Hart Flled' JuIy 1969 Attorney-Fulwider, Patton, Rieber, Lee & Utecht [2i 1 Appl. No.: 870,957

Related US. Application Data Division of Ser. No. 578,275, Sept. 9, 1966, abandoned.

US. Cl ..60/258, 60/3974 A, 60/265 lnt. Cl ..F02k 9/02 Field of Search ..60/258, 265, 39.74

References Cited UNITED STATES PATENTS 2/1954 Goddard ..60/258 [57] ABSTRACT Apparatus for fluid distribution and injection within a propulsive device, wherein a plurality of preformed laminations are stacked and bonded together to form an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, the interior surface having discharge pores communicating with corresponding fluid discharge channels within the shell.

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SHEET 7 [1F 7 INVENTOR. Evy/140 4 1 5400 %7'rae Ms v5 FLUID DISTRIBUTION AND INJECTION SYSTEMS This is a division of application Ser. No. 578,275, filed Sept. 9, 1966, now abandoned.

This invention relates generally to improvements in fluid distribution and injection systems of the type utilized in propulsive devices such as rockets and the like and, more particularly, to new and improved propellant distribution and injection systems for propulsive devices, resulting in improved economy, compactness, reliability, versatility, durability, and thermodynamic performance.

In the field of liquid fuel rocket type propulsion, it has been the general practice to employ injectors of various types to inject liquid propellants into a combustion chamber where the propellant burns and releases energy by chemical action to supply the motive power for propulsion.

The three principal types of propellants used in such engines are monopropellants, bipropellants, and hybrid propellants. Monopropellants are single liquids. Bipropellants consist of a fuel and an oxidizer, each carried separately within the flight vehicle and being brought together in the combustion chamber of the vehicle engine. Air breathing engines carry only fuel and use atmospheric oxygen for combustion. Hybrid propellants use a combination of liquid and solid materials.

The energy of liquid propellants is released in combustion reactions which also produce the working fluid for reaction propulsion. The liquids in a bipropellant system may ignite spontaneously upon contact (hypergolic liquids), or they may require an ignition device to raise them to the flash or ignition temperature (anergolic liquids). Combustion may be initiated with a spark, a heated wire, or an auxiliary hypergolic liquid. Monopropellant combustion, which is more properly a form of decomposition, can also be initiated by the catalytic action of an active surface or by a chemical compound in solution. Ignition of common hypergolic bi-propellants occurs typically in a period of 1-100 milliseconds after initial contact between the fuel and oxidizer. Sometimes catalytic quantities of various materials are utilized to decrease the ignition delay of certain specific bipropellants.

During engine operation, the combustion chamber contains a turbulent, substantially heterogenous, high temperature reaction mixture. The liquid propellants burn with droplets of various sizes in close proximity and travelling at relatively high velocities. During combustion, very high rates of heat release are encountered. Sometimes, unstable combustion due to nonuniform mixing and burning generates swirling gases which reach high velocities within the combustion chamber and excite the resonant frequencies of portions of the engine or other parts of the vehicle structure. Hence, many rocket engines use internal baffling to prevent such high frequency oscillation from occurring.

Overall engine performance, in contrast to theoretical propellant performance, is dependent upon effective design of the combustion chamber and injection system. Mixing and atomization are essential factors in injection of propellants into the combustion chamber. Injector and chamber design influence the flow pattern of both liquid and gases in the combustion chamber. One of these design factors in the characteristic chamber length referred to as L*, where L* is equal to the combustion chamber volume divided by the crosssectional area of the nozzle throat. In general, monopropellants require larger values of L* than bipropellants to provide the same performance in a rocket engine, due to the slower combustion exhibited by most monopropellants. For purposes of economy, it is desirable to keep L* to its lowest value which still provides relatively efficient combustion.

As previously indicated, in the field of liquid fuel rocket propulsion, one of the primary problems has been that encountered in connection with injection systems in the attempt to obtain optimum mixing and atomization of the propellants within the combustion chamber. Typically, a plurality of individual injectors would be used or, in order to provide greater discharge pore density, injectors were sometimes fabricated by means of drilling large numbers of very small holes in relatively thick metal plates. Each of the tiny drill holes in these injector plates was required to be angled precisely so that alternate streams of fuel and oxidizer, when forced through the injector, would impinge and mix in a manner to achieve good combustion in the rocket motor thrust chamber. Whenever holes were not drilled close enough or small enough, or whenever hole angles deviated from specified tolerances, the expense of fabricating the injectors escalated since the injector plates had to either be scrapped or painstakingly reworked at substantial cost. Moreover, even such injector plats often did not perform well from the standpoint of resultant combustion instability.

State of the art propellant injection systems generally tend to rely upon high propellant injection velocity to produce propellant atomization. Unfortunately, this requires high injection pressures and tends to limit the range of variation of injection pressure. Hence, the ability to throttle the thrust of an engine using such injection systems by means of control over injection pressure is extremely limited.

In connection with propellant injection problems, it is well known that the heat transfer coefficient of a rocket motor increases with decreasing combustion chamber size. State of the art injection systems, however, provide little or no cooling benefits. For this reason, small rocket engines usually become thermally limited, even when using conventional propellants. Where exotic propellants producing high flame temperatures, e.g., 8,000F. and more, are used, ablative combustion chambers are often incorporated into the engine design. However, it will be apparent that such ablative chambers, by virtue of their limited life, impose definite limitations over the duration of engine operation.

Hence, those concerned with the development of liquid fuel reaction type propulsion engines have long recognized the need for improved injection systems that provide satisfactory propellant dispersion, atomization and fuel/oxidizer mixing, as well as means for overcoming or minimizing the thermal problems encountered in small combustion chambers and with fuels producing high flame temperatures. The present invention fulfills these needs.

Accordingly, it is an object of the present invention to provide new and improved fluid distribution and injection systems which overcome the above and other disadvantages of the prior art.

Another object is to provide a new and improved fluid injection system which provides better atomizing of the injected fluid.

A further object of the invention is the provision of a new and improved fluid injection system which provides higher mixing efficiency.

Still another object is to provide a new and improved propellant injection system for propulsive devices providing enhanced design controllable fuel/oxidizer mixing characteristics.

Yet another object of the present invention is the provision of a new and improved fluid distribution and injection system for propulsive devices which enables reduced combustion chamber volume to produce the same thrust as previous devices.

A still further object is to provide a new and improved propellant injection system capable of high mixing efficiency at relatively low injection pressures.

Another object is to provide a new and improved propellant injection system capable of producing high mixing efficiency over a wide range of different propellant mass flow rates and, hence, capable of enhanced throttle control.

A further object of the invention is the provision of a new and improved propellant distribution and injection system for propulsive devices which enables more stable combustion and more uniform heat distribution.

Still another object is to provide a new and improved propellant distribution and injection system for propulsive devices which improves the thermal characteristics ofthe propulsive devices.

Yet another object of the present invention is the provision of a new and improved propellant distribution and injection system which is relatively economical to manufacture.

A still further object is to provide a new and improved fluid distribution and injection system in a propulsive device capable of simultaneously satisfying the cooling and combustion requirements of the propulsive device.

Another object is to provide a new and improved fluid distribution and injection system for propulsive devices characterized by building block capability over a relatively wide range of sizes.

Still another object is to provide new and improved propulsive devices integrally embodying fluid distribution and injection systems possessed of one or more of the afore-described advantages.

The above and other objects and advantages of the invention will become apparent from the following more detailed description, when taken in conjunction with the accompanying drawings ofillustrative embodiments thereof, and wherein:

FIG. I is a perspective view of a rocket engine embodying the fluid distribution and injection concepts of the present invention;

FIG. 2 is an exploded perspective view of the rocket engine shown in FIG. 1, a portion being broken out in section to illustrate structural details;

FIG. 3 is an enlarged plan view of the uppermost and lowermost lamination plates in the central plate stack shown in FIG. 2;

FIG. 4 is an enlarged, fragmentary perspective view of the area 4 in FIG. 3;

FIG. 5 is an enlarged, bottom plan view of the topmost end plate shown in FIG. 2;

FIG. 6 is a sectional view, taken along the line 15- 15 in FIG. 14;

FIG. 7 is an enlarged, sectional view, taken along the line 16-16 through the second plate from the top in FIG. 11;

FIG. 8 is an enlarged, bottom plan view of the second plate from the bottom in FIG. 1 1;

FIG. 9 is a sectional view, taken along the line l8-- 18 in FIG. 17;

FIG. 10 is an enlarged, sectional view, taken along the line 1919 through the lowermost end plate in FIG. 11;

FIG. 1 1 is an enlarged, plan view of the intermediate lamination plates in the central plate stack of FIG. 11;

FIGS. 12, 13, 14 and 15 are enlarged, sectional views through a stack of lamination plates of the type shown in FIG. 11, these views being taken substantially along the lines 12-12, 13-13, 14-14 and 15-15, respectively in FIG. 11.

FIG. 16 is a schematic flow diagram of that portion of the fuel distribution system for the rocket engine shown in FIG. 1 which accomplishes cooling of the throat region of the engine;

FIG. 17 is a flow diagram of the fuel and oxidizer distribution and injection system for accomplishing combustion within the rocket engine of FIG. 1;

FIG. 18 is a side view of a toroidal rocket engine embodying the present invention;

FIG. 19 is a left end elevational view taken in the direction of the arrow 19 in FIG. 18,

FIG. 20 is an enlarged sectional view, taken along the line 20-20 in FIG. 19, and illustrates the oxidizer distribution and injection wedges of the toroidal engine;

FIG. 21 is a left end view, taken in the direction of the arrow 21 in FIG. 20 and illustrates the profile of a single composite oxidizer wedge;

FIG. 22 is an enlarged, partial, sectional view, taken along the line 22-22 in FIG. 19, and illustrates the oxidizer and fuel distribution and injection plates between the oxidizer and fuel wedges of the toroidal engine shown in FIG. 18;

FIG. 23 is a left end view, taken in the direction of the arrow 23 in FIG. 22, and illustrates the profile of a small composite group of oxidizer and fuel lamination plates;

FIG. 24 is an enlarged, partial sectional view, taken along the line 24-24 in FIG. 19, and illustrates a fuel distribution and injection wedge;

FIG. 25 is a left end view, taken in the direction of the arrow 25 in FIG. 24 and illustrates the profile of a single composite fuel wedge.

Briefly, and in general terms, the present invention involves a fluid distribution and injection system wherein a plurality of preformed distribution and injection Iaminations are united together in a stacked arrangement to build up either a unitized fluid injector alone or a complete rocket engine integrally embodying a propellant distribution and injection system, all such devices making use of laminar controlled pore distribution and injection arrangements.

The term controlled pore is used generically herein to designate laminar distribution and injection systems wherein fluid distribution, manifolding, the number of discharge orifices or pores, the size, shape, density, spacing and array of the manifolding and orifices, as well as the injection angle and L/D ratio (ratio of discharge channel length to twice the effective hydraulic radius of the discharge orifice) of each orifice are all precisely controlled parameters to impart desired performance characteristics.

In the injector 200 of Flg. 9, all of the distribution apertures 208-211 of all of the injector plates of every type are in registry and define four vertically extending manifolds 208a-21la, respectively. The manifolds 209a and 211a are in fluid communication with each other by virtue of the manifold 220a, whereas the manifolds 208a and 210a are in fluid communication with each other via the manifold 222a. Hence, alternate vertical manifolds through the injector plates convey alternate fluids introduced via the conduits 205 and 206. It will be apparent, therefore, that the arrangement of converging discharge channels 213-218 for the injector plates 201, 202 and 203 in FIGS. 6, 7 and 8, respectively, is such that each discharge channel of a convergent pair sprays a different fluid than the other discharge channel of the same convergent pair. In this manner, the injector 200 provides a like on unlike fluid injection pattern.

Except for the difference in spraying format, i.e., like on unlike discharge vs. like on like discharge, the injector 200 of FIGS. 6-9 otherwise performs and possesses all of the advantages of the injector 100 illustrated in FIGS. 1-5.

Referring now to the drawings, and particularly to FIGS. l-l7 thereof, there is shown a new and improved rocket engine construction integrally embodying a fluid distribution and wraparound controlled pore injection system utilizing the same principles and laminar fabrication techniques previously set forth in the aforementioned divisional application Ser. No. 578,275, the disclosure of the latter application being specifically incorporated by reference in this application.

As best observed in FIGS. 1 and 2, a rocket engine 300, in accordance with the present invention, includes a plurality of stacked metal laminations for forming the combustion chamber and nozzle throat of the engine and distributing and injecting propellants introduced to the engine via a hollow fuel conduit 302 at the top of the engine and a hollow oxidizer conduit 303 at the bottom of the engine. The main body portion of the engine 300 is comprised of a stack of distribution and injector plates 305 (hereinafter referred to as injector plats) which are held between a pair of fuel distribution end plates 307, 308 at the top of the plate stack and a pair of oxidizer distribution end plates 310, 311 at the bottom of the plate stack. All of the plates are typically of copperclad stainless steel, the injector plates 305 being very thin, typically 0.007-0.0l0 inches in thickness with a copper coating of 0.0005 inches or less, the end plates being substantially thicker insofar as the base metal is concerned to withstand the intense heat released by combustion and to provide manifold channels of sufficient size for proper fluid distribution.

The manner in which the plates are stacked and bonded together under heat and pressure is the same as that previously described for the injectors 100 and 200 in the aforementioned divisional application Ser. No. 578,275. Hence, while the individual laminations have been shown, for purposes of illustration, in the engine 300 of Flg. 1, it will be apparent that the bonding process eliminates the distinct separate lamination character of the original plate stack and bonds all of the plats into'an integral unit having most of the properties of the stainless steel base metal.

It will also be apparent from FIGS. 1 and 2 that the number of injector plates 305 and, hence, the size of the plate stack, can be varied to build rocket engines over a wide range of different size and thrust output. Hence, the laminar fabrication and bonding technique enables building block capability for using the same parts to construct a wide range of different engines.

As best observed in FIG. 2, the stack of superposed injector plates 305 provides a wraparound array of discharge orifices 313 distributed over a substantial portion of the wall defining a combustion chamber 315 for the rocket engine. The injection pattern provided by the distribution of discharge orifices can be varied by the designer to simultaneously serve the cooling and combustion requirements of the rocket engine. In essence, therefore, the controlled pore discharge orifices 313 enable fluids to be injected into the combustion chamber 315 at any desired locations within the combustion chamber to best meet the needs of the engine for efficient cooling and combustion. If desired, the injection pattern may also extend into the throat region of the engine. As a result of this distributed injection pattern, not only is combustion efficiency increased, but the heat transfer rate to the chamber walls is also greatly reduced. Hence, smaller rocket engines can be provided which do not become thermally limited even when employing propellants having very high flame temperatures, such as nitrogen tetroxide/hydrazine or fluorine/hydrazine.

As observed in FIGS. 1-3, the provision of a venturi nozzle throat and the distributed arrangement of discharge orifices 313 integrally embodied within the engine 300 typically results in an engine horizontal cross-section bearing a resemblance to the small case Greek letter Omega and, hence, the engine 300 is referred to as an Omega motor. As shown in FIG. 3, the venturi throat of the Omega motor typically has a 30 convergent entrance from the combustion chamber 315 and a 15 divergent exit.

Referring now more particularly to FIGS. 2, 3 and 11 of the drawings, it is observed that the stack of injector plates 305 includes plates of two types 305a (FIG. 3) and 305b (FIG. 11). The injector plates 305a are located at the upper and lower ends of the plate stack, and typically only two or three of the plates 305a are located at each end of the stack. All of the remaining injector plates in the stack, and hence the great majority of the injector plates, are of the injector plate type 305b.

A comparison of FIGS. 3 and 11 indicates that the only difference between the injector plates 305a and 505b is that the plate 305a provides a pair of divergent fluid discharge channels 317 only from alternate fluid distribution apertures 318, whereas the injector plate 305b provides discharge channels 317 in divergent pairs from every distribution aperture 318. As will hereinafter become apparent, alternate distribution apertures 318 convey alternate fluids. Hence, the injector plate 305a will spray only a single fluid through the discharge channels 317 into the combustion chamber 315, whereas the injector plate 305b will spray two different fluids into the combustion chamber in pairs of impinging streams.

The single fluid injected by the plates 305a at each end of the plate stack is utilized as an evaporating cooling film to provide thermal protection for each of the distribution end plates 308 and 311, whereas the bulk of the injector plates 305b inject propellants into the combustion chamber 315 in a like on unlike impingement scheme such as that previously described in connection with the injector 200.

The fluid distribution apertures 318 and discharge channels 317 are best observed in the enlarged view shown in FIG. 4. As in the case of the distribution apertures and discharge channels of the injectors 100 and 200, each distribution aperture 318 extends all the way through the injector plate 305, and the discharge channels 317 are shallow channels of rectangular cross-section having a typical exit width of 0.01 inches and a typical exit height of 0.002 inches. The non-circular discharge orifice cross-section and angled discharge channel both contribute to the instability of the discharge stream which, as previously indicated, enhances the degree of atomization in the resulting spray.

When all of the injector plates 305 are stacked and bonded together to form an integral unit, the distribution apertures 318 of each injector plate are in registry with the corresponding distribution aperture of every other injector plate. The walls defining the distribution apertures 318 cooperate together to define a plurality of fluid distribution manifolds 318a, each manifold 318a being in fluid communication with a pair of discharge channels 317 in each injector plate 305b (FIGS. 2, 11 and 12), and alternate manifolds 3180 being in fluid communication with a pair of discharge channels 317 in each injector plate 305a (FIGS. 2 and 3).

As best observed in FIGS. 3 and 11, the fuel distribution aperture 318 on each side of the Omega shaped injector plates 305a, 305b nearest the venturi throat of the engine 300 is provided with a more divergent discharge channel 317a than the normal discharge channels 317. In this regard, the discharge channels 317a are angled to discharge fuel in an atomized spray nearly tangential to a wall of the combustion chamber 315 at the entrance to the venturi throat. The latter arrangement thus utilizes a small amount of fuel to evaporatively cool the venturi throat and thereby minimize throat errosion due to the high temperatures and heat transfer rates encountered during the combustion process.

Referring now to FIGS. 2, 3, 11 and 13-15, each of the injector plates 305a, 305b is provided with a pair of distribution apertures 320 on the entrance side of the venturi throat, symmetrically disposed on opposite sides of the combustion chamber 315 adjacent the last distribution apertures 318. Similarly, a pair of distribution apertures 322 are provided on the exit side of the venturi throat adjacent the open end of the engine. Each of the distribution apertures 320, 322 extends all the way through each injector plate 305, with shallow channels 323 extending between the pair of apertures 320, 322 on each side of the engine closely adjacent the throat defining interior surface of the combustion chamber wall.

As in the case of the distribution aperture 318 (FIG. 12), when the injector plates 305 are bonded together, the superposed apertures 320 define a pair of symmetrically disposed, vertically extending manifolds 320a (FIG. 14) and the distribution apertures 322 of each injector plate cooperate to define a pair of vertically extending fluid distribution manifolds 322a (FIG. 15) on the exit side of the venturi throat. The flow channels 323 in each injector plate 305 communicate between the manifolds 320a and 322a on each side of the engine and thereby provide a plurality of fluid flow cooling channels for further protecting the venturi throat section of the engine against the thermal effects of high temperature combustion.

The manner in which appropriate fluids are introduced from external sources (not shown) to the various distribution manifold channels 318a, 320a and 3220 of the engine by means of the distribution end plates 307, 308 and 310, 31 1 is next described.

Referring now to FIGS. 2 and 5-7 of the drawings, the fuel conduit 302 penetrates the outer face of the end plate 307 and is in fluid communication with a manifold channel 325 formed in the inner face of the end plate 307. The manifold channel 325 extends substantially across the width of the engine adjacent the open end of the engine and is adapted to mate with a corresponding manifold channel 326 in the end plate 308. Similarly, the inner face of the end plate 307 is provided with a second manifold channel 328 adapted to mate with a corresponding manifold channel 329 in the upper face of the end plate 308.

The inner face of the end plate 307 is also provided with a curved manifold channel 331 adapted to mate with a corresponding manifold channel 332 in the upper face of the end plate 308 to cooperatively define, upon bonding of the plates 307, 308 together, a manifold of sufficient extent around the perimeter of the engine 300 to overlie all of the distribution apertures 318 designated to receive fuel. In this regard, the upper face of the end plate 308 is provided with a plurality of sprues or flow channels 334 which interconnect the manifolds defined by the largerchannels 326, 329 and 332. Hence, fuel introduced via the conduit 302 to the manifold defined by the pair of mating manifold channels 325, 326 is introduced to all of the manifolds defined by abutment of the end plates 307, 308 together.

A plurality of apertures 336 extend through the end plate 308 and are located within the manifold channel 332 in positions such that the apertures 336 are in registry with alternate distribution apertures 318 in the plate stack 305. Similarly, apertures 338 are provided at opposite ends of the manifold channel 329 in registry with the apertures 320 of the plate stack 305, whereas apertures 340 are provided at opposite ends of the manifold channel 326 in registry with the apertures 322 of the plate stack 305. In this way, the end plates 307, 308 introduce fuel to alternate manifolds 318a and to the manifolds 320a and 322a formed in the plate stack.

The structure of the bottom end plates 310, 311 is similar to that of the top end plates 307, 308, except that provision is made for the introduction of oxidizer to the alternate manifolds 3180 which are not receiving fuel.

Referring to FIGS. 2 and 8-10 of the drawings, the inside face of the end plate 310 includes a pair of manifold channels 342, 344 of the same size and in precisely the same locations as the manifold channels 325 328, respectively, of the end plate 307 previously described. The manifold channels 342, 344 mate with a pair of corresponding manifold channels 346, 348 in the bottom face of the end plate 311 to define a pair of fuel manifolds. The latter fuel manifolds are interconnected by means of a plurality of flow channels 350 formed in the bottom face of the end plate 311 (FIG. 8).

A pair of apertures 340a are provided at opposite ends of the manifold channel 346 in the end plate 311, the apertures 340a corresponding to the apertures 340 in the end plate 308 and being in registry with the manifolds 322a formed in the injector plate stack 305. Similarly, apertures 338a are provided at opposite ends of the manifold channel 348 and these apertures are in registry with the manifolds 320a in the plate stack 305 in the same manner as the apertured 338 in the end plate 308.

Hence, a pair of fluid flow rings are defined by the end plates and injector plates for the flow of fuel introduced via the conduit 302. One such ring is located on the exit side of the venturi throat and is defined by horizontal manifold channels 325, 326 at the top of the engine 300, vertical manifolds 322a along both sides of the engine, and horizontal manifold channels 342, 346 at the bottom of the engine. The second fuel flow ring is located on the entrance side of the venturi throat and is defined by horizontal manifold channels 328, 329 at the top of the engine, manifolds 320a extending vertically along both sides of the engine, and manifold channels 344, 348 extending horizontally at the bottom of the engine.

As observed in FIGS. l3l5, the flow channels 323 communicate between the manifolds 320a and 322a on each side of the engine 300 and, hence, extend between the pair of fuel flow rings. Similarly, the channels 334 in the end plate 308 (FIG. 2) and the channels 350 in the end plate 311 (FIG. 8) interconnect the two fuel flow rings at the top and bottom of the engine.

It will be apparent, therefore, that fuel introduced by the conduit 302 is not only fed to the distribution manifolds 318a to be injected into the combustion chamber 315 for combustion purposes, but a portion of the fuel is also directed in a flow pattern for effectively cooling the venturi throat portion of the engine 300. This flow pattern is illustrated in FIG. 16 of the drawings wherein the two fuel flow rings are designated A and B, respectively.

Referring now again to FIGS. 2 and 8-10 of the drawings, oxidizer is introduced via the conduit 303 which passes through the bottom face of the end plate 310 into fluid communication with a manifold channel 352 formed in the inner face of the end plate 310 and adapted to mate with a corresponding manifold channel 353 formed in the bottom face of the end plate 31 1.

Abutting faces of the end plates 310, 311 are also provided with matching, curved manifold channels 355, 356, respectively, which extend around the perimeter of the end plates a sufficient distance to overlie all of the manifolds 318a which are designated to receive oxidizer. In this connection, the end plate 311 is provided with a plurality of apertures 358 within the manifold channel 356 and in registry with those alternate manifolds 318a in the injector plate stack 305 designated to receive oxidizer. Oxidizer is conveyed from the manifold defined by the channels 352, 353 to the manifold defined by the channels 355, 356 by means of flow channels 360 formed in the bottom face of the end plate 311 and extending between the manifold channels 353, 356.

Hence, it will be apparent that, for combustion purposes, fuel is introduced by the upper end plate assembly 307, 308 to alternate distribution manifold 318a in the injector plate stack 305, whereas oxidizer is introduced by the lower end plate assembly 310, 311 and is fed to the remaining alternate manifolds 318a in the injector plate stack. This injection scheme for the fuel and oxidizer components of a bipropellant system, and typically for hypergolic propellants such as hydrazene (fuel)/nitrogen tetroxide (oxidizer) is illustrated by the flow diagram in FIG. 17 of the drawings.

The cooling features of the Omega motor accomplished by means of the distributed injection pattern in the walls of the combustion chamber and the internal regenerative cooling at the venturi throat enhance the ability of the motor to withstand the intense heat 7 produced during the combustion process. Hence, small motors which usually become thermally limited because of their small combustion chamber size can be fabricated in accordance with the present invention in such sizes and utilizing propellant combinations which heretofore would have proven prohibitive because of thermal considerations.

The Omega motor also has a distinct advantage with regard to heat soak-back. When propellant flow is terminated, a certain amount of thermal energy is stored in any conventional combustion chamber. This heat energy soaks back into the propellant feed system and tends to vaporize the propellant in the system upstream of the flow cutoff valve typically utilized in such systems. However, because the combustion chambers in motors constructed in accordance with the present invention operate at much lower temperatures, there is therefore much less heat energy available to be transferred to the cutoff valve and upstream stations. In addition, since the amount of heat energy contained in the combustion chamber is a function of its mass, and smaller chambers are possible with the Omega motor, it will be apparent that such smaller motors will also tend to store and emit less heat energy.

Referring now more particularly to FIGS. 18-25 of the drawings, there is shown a rocket engine 400, integrally embodying a fluid distribution and wraparound controlled pore injection system utilizing the same principles and laminar fabrication techniques previously set forth in connection with the rocket engine 300.

The engine 400 is an Omega motor similar to the engine 300, except that the engine 400 is a torroidal engine having a shape defined by rotating an injector plate 305 (FIGS. 3 and 11) about an axis outside the plate 305 and parallel to the axis of symmetry. This body of revolution defining the rocket engine 400 is essentially provided by substituting wedge-shaped plates 405 for the planar injector plates 305 of the engine 300. The engine 400 also utilizes different manifolding arrangement for introducing and distributing fluids to the injector plates 405 of the engine 400 than the end plates 307, 308 and 310, 311 used in the engine 300 previously discussed. In all other respects, the rocket engine 400 is fabricated in the same manner, involves the same inventive concepts, and possesses all of the structural and performance advantages of the engine 300. However, since the toroidal engine 400 is a body of revolution requiring no end plates, end plate thermal problems are eliminated. Therefore, the toroidal engine 400 can withstand higher combustion chamber pressures and can produce greater thrust per unit volume than the rocket engine 300.

As best observed in FIGS. 18 and 19, the rocket engine 400, in accordance with the present invention, includes a plurality wedge-shaped stacked metal laminations for distributing and injecting propellants introduced to the engine via a hollow, circular oxidizer manifold 402 at the rear of the engine and a hollow, circular fuel manifold 403 at the front of the engine. The main body portion of the engine 400 is comprised of sector shaped stacks of wedge-shaped distribution and injector plates 405 which are held between pairs of narrower plate stacks comprised of oxidizer distribution and injection plates 406 or fuel distribution and injection plates 407.

The plates 406 are used to make up what is hereinafter referred to as an oxidizer wedge, whereas the plates 407 are used to make up what is hereinafter referred to as a fuel wedge. The oxidizer wedges 406 and fuel wedges 407 alternate in proceeding around the axis of symmetry of the engine 400 and, hence, each stack of injector plates 405 is sandwiched between an oxidizer wedge 406 and a fuel wedge 407. All of the plates 405, 406, 407 are typically of copper clad stainless steel, and it will be apparent that the bonding process eliminates the distinct separate lamination character of the original plate stack during assembly to unite all of the plates into an integral unit.

All of the oxidizer wedges 406 are in fluid communication with the oxidizer manifold 402 via appropriate conduits 402a (FIGS. 18 and 20). Similarly, all of the fuel wedges 407 are in fluid communication with the fuel manifold 403 via conduits 403a (FIGS. 19 and 25).

As best observed in FIGS. 20, 22 and 24, the plates 405, 406 and 407 cooperate to provide a wraparound array of discharge orifices 413 over a substantial portion of the wall defining a combustion chamber 415 for the rocket engine 400. As in the case of the rocket engine 300, the injection pattern provided by the distribution of discharge orifices 413 can be varied by the designer to simultaneously serve the cooling and combustion requirements of the rocket engine 400.

The plates 405, 406, 407 except for their wedge shape (FIGS. 19, 21, 23 and 25) embody the same distribution, injection and cooling concepts and structure as the plate 305b of FIG. 11 previously described in connection with the engine 300. Hence, the reference numerals 417-423 in FIGS. 20, 22 and 24 for the engine 400 denote like structure corresponding to that indicated by the reference numerals 317-323, respectively, for the plate 305b in FIG. 1 1. Hwever, it will be apparent that, when all of the plates 405, 406 and 407 are stacked and bonded together to form an integral unit, the distribution apertures 418, 420 and 422 plate,

each plat, by virtue of the wedge shape of the plates, cooperate together to define a plurality of circular fluid distribution manifolds for the engine 400 instead of the linear manifolds formed by the corresponding apertures 318, 320, 322, respectively, for the engine 300.

As best observed in FIG. 20, each plate of the oxidizer wedge 406 includes an arc-shaped distribution aperture 425 in fluid communication with alternate apertures 418. When several of the plates 406 are stacked together, the walls defining the distribution apertures 425 cooperatively define a distribution manifold in fluid communication with the external oxidizer manifold 402 via the conduit 402a. In this connection, the spacing between the distribution apertures 418 and the outer perimeter of the plate 406 is greater than that for the plate 405 in FIG. 22, in order to provide sufficient space for the distribution aperture 425.

Referring now to FIG. 24, each of the plates defining the fuel wedge 407 is also provided with an arc-shaped distribution aperture 427 for building up a fuel distribution manifold in fluid communication with alternate distribution apertures 418 designated to receive fuel. The distribution aperture 427 is also in communication with the apertures 420 which receive fuel from the external fuel manifold 403 via conduit 403a, apertures 422 and flow channels 423.

The controlled pore techniques and apparatus of the present invention satisfy a long existing need in the propulsion art for distribution and injection systems which are economical, compact, durable, and having improved thermodynamic performance and throttle control characteristics.

It will be apparent from the foregoing that, while particular forms of the invention have been illustrated and described, various modifications can be made without departing from the spirit and scope of the invention. In this regard, while the specific embodiments of the invention described herein are directed to systems for distributing and injecting more than one fluid, such systems can readily be adapted for the discharge of only a single fluid, such as a monopropellant. Accordingly, it is not intended that the invention be limited, except as by the appended claims.

I claim:

1. A propulsive device, comprising:

a plurality of preformed, Omega-shaped, wedge laminations assembled in a stack and bonded together to provide an integral structural shell in the shape of a toroid, said shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of said preformed laminations defining spaced apart layers of small propellant discharge pores distributed over a substantial portion of said interior surface;

a plurality of fluid flow discharge channels formed within said shell, the number of said discharge channels being equal to the number of said discharge pores, each discharge channel terminating at one of said discharge pores in said injection surface, whereby each discharge pore is associated with a single discharge channel;

manifolding means defined by said laminations for distributing a first propellant fluid to some of said discharge channels and associated discharge pores and for distributing a second propellant fluid to others of said discharge channels and associated discharge pores; and

fluid distribution means formed by said laminations within said structural shell for utilizing at least a portion of one of said propellant fluids, prior to injection, for cooling said nozzle throat.

2. A propulsive device as set forth in claim 1, wherein said discharge pores are of non-circular shape, and said discharge channels are angled within said shell such that the fluid stream emitted from one of said discharge pores in a layer impinges upon the fluid stream emitted from an adjacent discharge pore in the same layer.

3. A propulsive device, comprising:

a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell, each of said laminations and said structural shell having an Omega cross-section, said shell having an interior surface defining a combustion chamber, adjacent pairs of preformed laminations defining spaced apart layers of small propellant discharge pores distributed over a substantial portion of said interior surface, said interior surface also defining a nozzle throat in addition to said combustion chamber.

4. A propulsive device, comprising:

a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surface;

a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore is associated with a single discharge channel, said discharge channels being angled within said shell such that the fluid stream emitted from one of said discharge pores in a layer impinges upon the fluid stream emitted from a second of said discharge pores in the same layer; and

manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores.

5. A propulsive device, comprising:

a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surface;

a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore 15 associated with a single discharge channel,

said manifold means distributing more than one fluid to different discharge channels of the same layer to provide a like on unlike fluid injection pattern; and

manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores.

6. A propulsive device, comprising:

a plurality of preformed wedge-shaped laminations assembled in a stack and bonded together to provide an integral structural shell which is toroidal in shape, said shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surface;

a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore is associated with a single discharge channel; and

manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores. 

1. A propulsive device, comprising: a plurality of preformed, Omega-shaped, wedge laminations assembled in a stack and bonded together to provide an integral structural shell in the shape of a toroid, said shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of said preformed laminations defining spaced apart layers of small propellant discharge pores distributed over a substantial portion of said interior surface; a plurality of fluid flow discharge channels formed within said shell, the number of said discharge channels being equal to the number of said discharge pores, each discharge channel terminating at one of said discharge pores in said injection surface, whereby each discharge pore is associated with a single discharge channel; manifolding means defined by said laminations for distributing a first propellant fluid to some of said discharge channels and associated discharge pores and for distributing a second propellant fluid to others of said discharge channels and associated discharge pores; and fluid distribution means formed by said laminations within said structural shell for utilizing at least a portion of one of said propellant fluids, prior to injection, for cooling said nozzle throat.
 2. A propulsive device as set forth in claim 1, wherein said discharge pores are of non-circular shape, and said discharge channels are angled within said shell such that the fluid stream emitted from one of said discharge pores in a layer impinges upon the fluid stream emitted from an adjacent discharge pore in the same layer.
 2. A propulsive device as set forth in claim 1, wherein said discharge pores are of non-circular shape, and said discharge channels are angled within said shell such that the fluid stream emitted from one of said discharge pores in a layer impinges upon the fluid stream emitted from an adjacent discharge pore in the same layer.
 3. A propulsive device, comprising: a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell, each of said laminations and said structural shell having an Omega cross-section, said shell having an interior surface defining a combustion chamber, adjacent pairs of preformed laminations defining spaced apart layers of small propellant discharge pores distributed over a substantial portion of said interior surface, said interior surface also defining a nozzle throat in addition to said combustion chamber.
 4. A propulsive device, comprising: a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surface; a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore is associated with a single discharge channel, said discharge channels being angled within said shell such that the fluid stream emitted from one of said discharge pores in a layer impinges upon the fluid stream emitted from a second of said discharge pores in the same layer; and manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores.
 5. A propulsive device, comprising: a plurality of preformed laminations assembled in a stack and bonded together to provide an integral structural shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surfaCe; a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore is associated with a single discharge channel, said manifold means distributing more than one fluid to different discharge channels of the same layer to provide a like on unlike fluid injection pattern; and manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores.
 6. A propulsive device, comprising: a plurality of preformed wedge-shaped laminations assembled in a stack and bonded together to provide an integral structural shell which is toroidal in shape, said shell having an interior surface defining a combustion chamber and a nozzle throat, adjacent pairs of preformed laminations defining spaced apart layers of small fluid discharge pores distributed over a substantial portion of said interior surface; a plurality of fluid flow discharge channels formed within said shell between adjacent pairs of preformed laminations, the number of discharge channels being equal to the number of discharge pores, each discharge channel terminating at one of said discharge pores, whereby each discharge pore is associated with a single discharge channel; and manifold means formed within said shell by said laminations for distributing fluid propellant to said discharge channels and associated discharge pores. 